Replacement of machined, cast, forged or stamped metallic structural elements such as aluminum wing ribs, in an aerospace industry example, with composite structures is becoming highly desirable for weight reduction and performance enhancement of structural assemblies. For the composite wing rib example, without limitation, ribs may be similar to built-up aluminum ribs, using separate webs, chords and stiffeners. All of the parts may be produced primarily from epoxy pre-impregnated composite tape which typically is placed manually on a predetermine shaped mandrel followed by curing within an autoclave.
The two fundamental approaches presently in the art are to assemble composite laminate onto a tooling form in a predetermined manner in its final shape on the cure mandrel, or assemble in a predetermined manner a flat charge of composite laminate, mechanically form it to shape, and then cure it.
Forming of composite elements which have variable contour with local areas having a contour as tight as 300 inch radios currently requires manual, layup on a shaped mandrel. At a typical rate of one pound per boor per employee, manual laminate assembly of the large structural elements is not art economical approach for full production applications. Automated lay-tip of composite structural elements directly on cure mandrels is possible with modified fiber placement machines. However, such an approach requires specific machine development with associated very high capital investment.
It is therefore desirable to assemble in a predetermined, manner flat composite laminate assemblies or charges and form them to shape. Flat charges are typically manufactured using composite tape automation methods and these machines are very cost effective at producing flat charges. Capability is readily available with current manufacturing processes to drape form flat charges to a channel shape on a flexible mandrel and then cute the mandrel. However, this process is suitable for large radius bends only. At radii of 1500 inches or less, there may be problems of causing buckles in the pre-preg composite laminate assembly.
It is also therefore desirable to provide a method and tooling for shaping and cure of flat charge layups of prepreg materials with single or multiple centers of curvature which precludes buckling of composite assembly.